The present invention relates to a blade for the rotary wings of an aircraft, in particular a helicopter, said blade making it possible simultaneously to reduce noise and improve the performance at high load, especially at takeoff and at moderate flight speed.
It is known that, both in hovering flight and in forward flight, the performance of a rotor of a rotary-wing aircraft, especially a helicopter, is limited by the following phenomena:
the shockwaves which develop on the suction face of the advancing blades during high-speed flights;
the stalling that results from the detachment of the boundary layer on the suction face of the retreating blades when there is a demand for lift in translational flight;
the interaction of the vortex generated by the previous blade with the following blade, which leads, during hovering flight, to a substantial dissipation of energy in two forms: induced power and profile drag power.
In addition to being responsible for loss of performance, the shocks and the blade-vortex interaction are also responsible for acoustic problems in the form of pulsating noise, caused by shock delocalization (high-speed flight) and by pulsating changes in lift when the marginal vortex directly strikes the blade (decent), respectively.
It has been found that the performance of a blade for the rotary wings of an aircraft depends, to a large extent, on parameters associated with the construction of the blade, such as:
a) the radial distribution of the blade area;
b) the sweepback of the blade tip;
c) the change in relative thickness of the profiles;
d) the distribution of the twist of the profiles; and
e) the blade tip droop.
The influence that the three main parameters a), b) and c) have on the performance and on the noise of a rotary-wing blade is explained in detail below.
a) Radial Distribution of the Blade Area
For a rotor of a rotary-wing aircraft whose elementary profiles or sections all work with the same coefficient of lift Cz, the linear lift varies as the local chord length L(r) and as the square of the local speed, which is directly proportional to the radius (radial position) r of the section. This means that the total lift of the blade varies proportionally with the mean chord {overscore (L)} defined by a square-law weighting of the radius r:       L    _    =                    ∫        R0        R            ⁢                        L          ⁡                      (            r            )                          ⁢                  r          2                ⁢                  ⅆ          r                                    ∫        R0        R            ⁢                        r          2                ⁢                  ⅆ          r                    
in which RO represents the radius r at the start of the blade section at the root end of the blade and R is the total radius of the blade.
It is common practice for the performance of blades of various shapes to be compared, by referring them to this mean chord {overscore (L)}.
Compared with a conventional blade of rectangular shape, calculations show, and experience confirms this, that a reduction in the chord at the outboard end of the blade (tapered shape) improves performance over a wide speed range, including in hovering flight. In translational flight, this improvement is explained essentially by the reduction in the drag of the profiles, which is due to the reduction in the profile drag due to the reduction in the chord at the tip. Shocks in this region are exerted on a smaller area while the central part of the blade, not subjected to the shocks, provides most of the lift with a maximum efficiency: the lift/drag ratio here is a maximum. At low speed and in hovering flight, the tapering of the tip improves the lift efficiency-that is to say makes it possible to reduce the power needed for a given lift force-according to a different mechanism: a more homogeneous induced velocity distribution is obtained over the entire rotor disk, preventing the load from being too concentrated at the tip. The distribution thus approaches the optimum distribution for the lift efficiency, which consists of a uniform induced velocity over the entire disk.
A second known advantage provided by the outboard taper of the blade is a certain reduction in the noise. On the one hand, the volume of air displaced by the high-speed passage of the tip is reduced as the square of the chord (for the same relative thickness of the profiles). This results in a reduction in some of the noise still present, that corresponding to the noise called xe2x80x9cmonopole sourcexe2x80x9d. On the other hand, the blade edge vortex, at the origin of the blade-vortex interaction noise, curls more slowly and the maximum velocity in the core of this vortex is lower the greater the distance between the maximum chord zone and the tip. This results in appreciable attenuation of the interaction noise, particularly during the decent phases of the aircraft.
However, the outboard tapering of the blade has the drawback of requiring an increase in the chord over the rest of the span, so as to maintain the constant mean chord {overscore (L)} and so as not to excessively increase the coefficient of lift Cz of the profiles. This increase in the chord may be significant because of the r2 weighting in the expression for {overscore (L)} (see above) and this results in the rotor being somewhat heavier. Nevertheless, the tapering on the blade tip side over a moderate length, of the order of 5 to 6% of the rotor radius, is a means commonly employed for improving the performance of the latter, generally in combination with sweepback of the blade tip, as illustrated in Patents FR-2,755,941, FR-2,689,852 and FR-2,617,118.
The tapering of the chord on the inboard side of the blade, that is to say on the side where it is attached to the hub, is a known means of limiting the drawback of an increase in mass and of improving the performance at high speed, that is to say above 300 km/h, since, under these conditions, this zone of the blade contributes little to the lift and greatly to the power consumed by the rotor (see Patents FR-2,755,941 and FR-2,689,852). However, this arrangement proves to be unfavorable in the case of the performance at moderate speed and in hovering flight since it tends to excessively reduce the load in the central zone of the rotor and to make the induced velocity distribution less uniform, thereby resulting in a reduction in the lift efficiency.
b) Offset of the Profiles in the Plane of the Rotor, with Part of the Blade Swept Back
In addition, in order to push back the threshold at which shockwaves appear and to limit their intensity, it is advantageous for the blade tip to be curved nearward (Patents FR-2,755,941 and FR-2,689,852 and Patent Application FR-97/16227) or else for it to have a double curvature, alternately toward the front and the rear (Patent FR-97/11230). The sweep angle xcex9, defined by the line of aerodynamic centers (approximately at the front quarter of the chord) and the feathering axis, reduces the effective Mach number and thus sweeping back the blade tip constitutes an effective means of reducing the unfavorable consequences of the compressibility of air, especially the appearance of shockwaves.
However, it is new known (Patent Application FR-97/16227) that the sweep angle must remain modest, typically less than 35xc2x0, so as to avoid the formation of a three-dimensional ram""s horn vortexxe2x80x94or apex vortexxe2x80x94similar to that observed on delta-type wings. This is because this type of very stable and concentrated vortex produces intense interactions with the following blades and therefore contributes to the noise of a helicopter as it descends.
Furthermore, the magnitude of the offset with respect to the feathering axis, and the span length of the zone in question, also limited by the torsional forces which result from the offset of the aerodynamic lift as well as of the center of gravity. A known means for limiting this unfavorable effect consists in shifting the profiles of the mid-part forward and those of the tip rearward in such a way that the blade remains balanced overall: see, in particular, Patents FR-2,755,941, FR-2,689,852 and FR-97/11230.
c) Change in the Relative Thickness of the Profiles
The relative thickness of a blade cross section is defined as the ratio of the absolute thickness e with respect to the chord length L of the profile that constitutes the contour of this section, i.e. e/L.
Blades whose shape is optimized exclusively for high speeds (Patent FR-2,755,941) or for noise reduction (Patent FR-97/11230) exhibit a variation in the relative thickness of the profiles which increases slowly from the tip toward the central part, with a thickness approximately equal to 7% at the tip and remaining less than 11% at the section located at 75% of the rotor radius. This arrangement, which is favorable to the high-speed performance with a moderate lift and allows noise reduction, is nevertheless subject to premature stalling whatever the speed, including in hovering flight. In general, this arrangement proves to be of low performance with strong lift.
It has in fact been discovered that the profiles offering the best compromise between efficiency (Cz/Cx ratio for moderate lift) and maximum lift (Czmax) for sections lying between 50% and 75% (Mach number between 0.3 and 0.6) have a relative thickness of between 12 and 14%.
It may therefore be seen that none of the above documents describes a blade structure entirely without drawbacks.
The object of the present invention is to remedy these drawbacks. It relates to a blade for a rotary wing, the geometry of which is optimally defined so as to guarantee the best performance for an aircraft, especially a helicopter, the lift of which is provided by such a wing, said aircraft flying within a moderate speed range, of between, for example, 0 (hovering flight) and 300 km/h, and being capable of flying and maneuvering with a high load in hot weather and at altitude, these conditions being characterized by a mean coefficient of lift Czm of the profiles which can reach a value of 0.7. This blade must furthermore generate limited noise, especially during landing phases.
For this purpose, according to the invention, the blade with a swept-back tip for the rotary wings of an aircraft, intended to form part of a rotor whose hub is linked to said blade, which blade is capable of being driven in rotation about the axis of said hub, said blade having a leading edge and a trailing edge, and being formed from successive elementary cross sections identified by the distance r which separates each of them from the rotation axis of said hub, and each having a defined chord profile and an aerodynamic center whose offset with respect to the feathering axis, orthogonal to each of said sections, determines the sweepback of said blade, is remarkable in that said blade being subdivided along its longitudinal extent into three zones, namely a first zone extending from the inboard end RO of the blade to a section R1 located at approximately 90% of the total length of the blade, a second zone extending from the section R1 to a section R2 located at approximately 95% of the total length of the blade and a third zone extending from the section R2 to the free outboard end R of the blade:
the length of the chord L is a maximum and approximately constant in said first zone, decreases linearly in said second zone and decreases according to a parabolic function in said third zone while respecting the continuity of the rate of variation of the chord at the common limit with the second zone; and
the offset Yxe2x80x2f of the aerodynamic center with respect to the feathering axis is approximately zero in said first and second zones and decreases according to a parabolic function in said third zone, while respecting the continuity of the sweep angle at the common limit with the second zone. This decrease is furthermore chosen so that the trailing edge remains straight and continuous along said second and third zones.
Thus, the blade geometry as defined makes it possible to guarantee the optimum performance for a rotary-wing aircraft, especially a helicopter, the lift of which is provided by a wing consisting of such blades, said aircraft flying within a moderate speed range of between 0 and 300 km/h for example, with a mean coefficient of lift of the profiles Czm which can reach a value of 0.7.
In addition, by virtue of the invention, said blade produces limited noise, especially during landing phases (reduction in blade-vortex interaction noise).
With regard to the variation in chord length, the invention provides for quite a large taper on the tip side (uniformity of the induced velocity in hovering flight and reduction in the volume at the tip in order to minimize noise of monopolar origin) without there having to be a taper on the attachment side (uniformity of induced velocity and maximum torsional rigidity).
Thus, the start of the taper is further from the tip than in the case of the blade disclosed in Patent FR-2,617,118 and is located more or less at the same position as in Patent FR-2,755,941 or FR-2,689,852, but with a chord which remains constant toward the blade root.
The recommended chord variation, as well as the maximum and minimum limiting curves between which the abovementioned advantages may be obtained, will be specified below.
Moreover, since the aerodynamic center is defined here as the point on each section lying at the front quarter between the leading edge and the trailing edge, the offset Yxe2x80x2f is the distance, in the direction of the chord, between the feathering axis and said center, counted positively when the displacement of the section takes place toward the leading edge. The sweep angle xcex9 is defined as the angle between the tangent to the curve joining the centers of the sections and the feathering axis, seen from above. The sweep is directed rearward from the tip side of the blade. The angle xcex9 is calculated directly from the law of variation of Yxe2x80x2f:xcex9(r)=arctan(dYxe2x80x2f/dr).
Given the envisaged speed range, the favorable effect of the sweep angle reducing the intensity of the shocks is fully obtained for values of the sweep angle xcex9 not exceeding 33xc2x0, this maximum value being reached only at the blade tip. This value, which is markedly smaller than those recommended in Patents FR-2,755,941, FR-2,689,852 and FR-2,617,118, is intended to avoid the formation of a ram""s horn leading-edge vortex.
In the present invention, the angle xcex9 is limited by virtue of the very gradual reduction in the chord, combined with a straight trailing edge over the entire blade part lying between the widest chord and the tip.
Another favorable consequence of this novel arrangement is that the rearward offset of the aerodynamic center at the tip, with respect to the feathering axis, remains particularly small so that the torsional forces and the forces to which the pitch control system is subjected remain very moderate (without it being necessary to produce balancing by an offset in the opposite direction of the internal part of the blade).
The division of the blade into three zones, as described above, applies in the same way to the law of variation of the offset Yxe2x80x2f and the sweep angle xcex9 which stems therefrom. As indicated above, in the first two zones, the offset is approximately zero and, in the third zone, it decreases parabolically while respecting the continuity of the angle xcex9 at the common limit with the second zone.
The recommended variation (law of variation) of the offset Yxe2x80x2f, together with the maximum and minimum bounding curves between which the aforementioned advantages may be obtained, will be specified below.
Within the context of the present invention, the best blade performance is obtained by combining the variation of the chord length L and that of the offset Yxe2x80x2f of the aerodynamic center, as defined above, with a particular variation, and in accordance with the invention, of the relative thickness of the blade.
For this purpose, advantageously, said blade being, in addition, subdivided along its longitudinal extent into three additional zones, namely a first additional zone extending from the inboard end RO of the blade to a section R3 located at approximately 50% of the total length of the blade, a second additional zone extending from the section R3 to a section R4 located at approximately 75% of the total length of the blade and a third additional zone extending from the section R4 to the free outboard end R of the blade, the relative thickness of the blade, which corresponds to the ratio and the absolute thickness e to the chord length L:
decreases approximately linearly with an approximate slope of xe2x88x920.12% of thickness per % of span in said first additional zone so as to end up with a relative thickness approximately equal to 12% in the section R3;
remains constant and approximately equal to 12% in said second additional zone; and
decreases approximately linearly in said third additional zone so as to reach a value approximately equal to 7% at the free outboard end of the blade.
The recommended variation in the relative thickness, together with the maximum and minimum limiting curves between which the abovementioned advantages may be obtained, will be specified below.
Moreover, in one particular embodiment, the vertical displacement Zv from the center of twist with respect to the plane of zero lift of the blade is such that the center of twist remains approximately in said plane in the aforementioned first and second zones (relative to the chord length and to the offset Yxe2x80x2f), and in the third zone:
Zv(r/R)/R=0.0905 x2 with x=(rxe2x88x92R2)/(Rxe2x88x92R2).
This variation in the vertical offset or displacement Zv makes it possible to improve the lift efficiency at takeoff.
However, it is also conceivable within the context of the invention to apply no vertical offset so as to limit the vibration excitation in translational flight, renouncing the advantage provided in terms of lift efficiency.
Furthermore, advantageously, the blade has a linear aerodynamic twist with a total amplitude of approximately xe2x88x9210xc2x0 between the center of the rotor and the free outboard end R of the blade. In accordance with common practice, the twist is counted negatively when the leading edge of the outer sections is reduced with respect to that of the sections lying closer to the center. In order to obtain the geometrical setting of each section, counted with respect to the reference chord, it is necessary to add (algebraically) the effect of the zero lift of the profile in question on the aerodynamic twist.